The present invention relates generally to gas turbine engines, and, more specifically, to turbine nozzles therein.
In a gas turbine engine air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases. Energy is extracted from the hot gases in several turbine stages following the combustor for powering the compressor and an additional output shaft for turbofan aircraft engines, or marine and industrial engine applications.
The turbine stages include stationary turbine nozzles which direct the combustion gases into corresponding rows of turbine rotor blades extending radially outwardly from the perimeter of supporting rotor disks. The first stage turbine nozzle is disposed at the outlet of the combustor and includes a row of hollow vanes supported in radially inner and outer bands.
The hot gases discharged from the combustor effect substantial thermal loads in the turbine nozzle, and cause substantial thermal expansion and contraction thereof during the operating cycle of the engine. In order to reduce thermal stresses in the nozzle, the nozzle is typically segmented circumferentially in vane doublets including two vanes mounted in corresponding inner and outer band segments. The segments have end faces or axial splitlines in which spline seals are used to connect the row of doublets in a complete annular assembly.
The spline seals seal the end faces of the nozzle segments from the hot combustion gases flowing through the turbine flowpath, as well as from the cooling air being circulated outside the outer band and inside the inner band for cooling thereof. And, the outer and inner bands adjoin the outlet end of the combustor on the forward side thereof, and adjoin the inlet end of the first stage turbine rotor blades on the aft side of the nozzle, with suitable seals therebetween.
The individual nozzle vanes include internal cooling circuits or cavities through which cooling air bled from the compressor is circulated for providing internal cooling of the nozzle vanes during operation. And, the nozzle vanes typically include various rows of film cooling holes extending through the pressure and suction sidewalls thereof for discharging the spent internal cooling air in corresponding films of air along the external surface of the vanes for providing thermal insulation from the hot combustion gases flowing thereover during operation.
The individual nozzle vanes have aerodynamic profiles, with the pressure side being generally concave and the opposite suction side being generally convex between the leading edge of the vane to the relatively thin trailing edge of the vane. The velocity and pressure distributions of the combustion gases correspondingly vary over the opposite sides of each vane for maximizing performance of the turbine in extracting energy from the combustion gases to power and rotate the turbine blades supported on the rotor disk.
In one exemplary gas turbine engine, experience has uncovered oxidation damage to the inner band of the first stage turbine nozzle which limits its useful life. The damage is localized to one corner of the inner band at the aft end thereof along the suction side of the nozzle vane. This oxidation may be due to hot gas ingestion between the inner band end faces at the aft end of the splitline due to random variations in differential pressure between the hot combustion gases on the turbine flowpath side of the inner band, and the cooling air provided on the opposite cooling side of the inner band.
Accordingly, it is desired to provide a turbine nozzle having improved cooling of the inner band to address this local oxidation problem and increase the useful life of the turbine nozzle.